Pitch control system for aircraft



Aug. 16, 1960 c. A. RAFFERTY PITCH CONTROL SYSTEM FOR AIRCRAFT Filed March 17, 1953 INVETOR. CHRISTOPHER A. RAF'FERTY United States Patent PITCH CONTRGL SYSTEM FOR AIRCRAFT Christopher A. Rafierty, Brooklyn, N.Y., assignor to Bendix Aviation Corporation, Teterboro, N.J., a corporation of Delaware Filed Mar. 17,1953, Ser. No. 342,781

14 Claims. (Cl. 244- 77) This invention relates generally to control systems and more particularly to automatic control systems for aircraft.

Conventional practice is to control the attitude of an aircraft by a gyroscope. The disadvantage of this is that While the weight of the aircraft is substantially constant, the lift developed by the wings at any instant depends in part upon the relative wind; relative wind being defined as the direction of air flow with respect to an airfoil. When the speed of the relative air stream is in creased, the craft gains in altitude; conversely, when the speed of the air stream is decreased, the craft loses altitude. Within limits, by controlling the attitude of the craft with respect to the relative air, i.e., the angle of the aircraft with respect to the relative wind greater efficiency can be achieved since the craft could be maintained at a constant level despite changes in the speed of the relative wind. The angle of an airfoil of the aircraft with respect to the relative wind is termed the angle of attack.

An object of the present invention, therefore, is to provide a novel means by which the angle of attack of a craft may be controlled With respect to the speed of the relative air stream.

Another object is to provide a novel means for controlling the rate of climb of aircraft without afiecting its airspeed and Without introducing large transient disturbances in airspeed.

correlating the pitch attitude of an aircraft with the thrust of its power plant and the rate of climb.

Another object is to provide a novel means for indicating the angle of attack of an aircraft for quiescent air, i.e., with no vertical component.

A further object is to provide a novel apparatus for comparing the lift coelficient of the craft with the lift coeflicient required for level flight.

A still further object is to provide a novel apparatus for indicating the angle of the flight path with respect to the horizon. i

' The above and further objects and novel features of the invention Will appear more fully from the following detailed description when the same is read in connection with the accompanying drawing. It is to be expressly understood, however, that the drawing is not intended as a definition of the limits of the invention but is for the purpose of illustration only.

In the single sheet of drawing, Figure 1 illustrates a complete schematic Wiring diagram of the pitch channel of an automatic pilot system embodying the present invention.

Figure 2 illustrates the vectorial relationship which occurs in a resolver in Figure 1.

In U.S. Patent No. 2,625,348, there was described an automatic control system for aircraft. The accompanying drawing illustrates the pitch control channelof an "2,949,257. Patented-Aug. 16, 1960 '2 automatic pilot system, which may be the automatic pilot system described in the aforementioned patent, as modified in accordance with the present invention.

Pitch control is attained in the above patent by having a servomotor 10 through a gear train 12 move elevators 14. To prevent hunting, the gear train also operates a' follow-up device 16 and a rate generator 18. The follow-up device develops a signal corresponding to the displacement of the elevator surface from its normal streamlined position, and the rate generator develops a signal corresponding to the speed of rotation of the servomotor.

Servomotor 10 operatesin response to a combination of signals. In this chain of signals, the combined rate generator and follow-up signal from potentiometer 20 is added to a signal which corresponds to the rate of turn of the craft about its pitch axis and is developed by an inductive device 21 which is actuated by a pitch rate gyroscope 22. These signals are added with a signal from an inductive device 23 which is actuated by a vertical gyroscope 24, the latter signal corresponding to the pitch attitude of the craft. To this combination is further added the signals developed by a novel angle of attack computer 26 and a flight path angle computer 28. This combination of signals after amplification by an amplifier 30 operates motor 10 to maintain the craft at the desired pitch attitude.

Neglecting the follow-up, rate generator, and pitch rate gyro signals which are for damping purposes, the pitch attitude of the aircraft is governed by a summation of the signals from the angle of attack computer 26, the

- flight path angle computer 28, and the pitch take-off 23' on the vertical gyro 24. The actual pitch attitude of the craft is a summation of the angle of attack and the flight path angle.

When the aircraft is in level flight, the lift L developed by the wings is equal to the weight W'of the craft or This lift may be calculated from the equation 2 2 L= %/SC where:

p=air density in slugs per cubic foot V=true airspeed in feet per second S=wing area in square feet C =coeflicient of lift, which is dependent on the wing contour and includes the mathematical calculations that make it possible to determine the lift by multiplying the density, area, and velocity factors for the particular condition by this coefficient.

The airspeed may be measured by the dynamic pressure which may be determined by measuring the difference between the impact and static pressures in a Pitot tube or airspeed head. Dynamic pressure may be defined mathematically as:

Substituting the valve q of Equation 3 for its equivalent in Equation 2 7 Substituting the value of L of Equation 1 in Equation 4 V 2,949,257 I l t 3 Now, the lift coefiicient at any angle of attack is:

(10' CL: no-lwhere: x

C =lift coeificient at zero angle of attack ot=angle of attack =slope of the lift coeificient curve versus the V angle of attack at the speed of flight.

The quantities C and dd at conventional aircraft speeds are generally constant.

From Equations 6 and 7 can be written (8) W/qs= m+ Solving for on W 1 1.0 9) ,a=( -E) T do: da

Let at; be the angle of attack measured from zero lift 1 (10) ar'=(q g fi d 11 W/S=a q(% The total angle of attack at any time is h 12 E da The novel device for computing the angle of attack is comprised generally of five potentiometers 51, 52, 53, 54, 55. Potentiometer 55 is a trim potentiometer while potentiometers 51,52, 53 and 54 supply functions corresponding to the terms q, a

and W/S, respectively, of Equation 11. Potentiometers 51, 52, and 53 supply a signal corresponding to the right hand side of Equation 11, and potentiometer 54 supplies a signal corresponding to the left hand side of the equation. While potentiometers 51 and 54 are energized from a suitable source of alternating current, the energization of potentiometer 51 is opposite in phase to the energization of potentiometer 54. The wipers of potentiometer 53 and 54 may be positioned manually by suitable means such as knobs 56 and 58, the wiper of potentiometer 52 may be moved by a motor 60, and the wipers of potentiometers 51 and 55 by an airspeed responsive device 62.

Airspeed responsive device 62 may be of. a. conventional type, consisting of a diaphragm 63 having one side subjected by way of a tube 65 to static air pressure and the other side subjected by Wayof a tube 67 to impact pressure. A gear train 69 transmits the movement of diaphragm 63 in response to a difierence in pressure to a shaft 71.

Considering now the right hand portion of Equation 11, the signal corresponding to the dynamic pressure q is derived from potentiometer 51 whose wiper is fixed to shaft 71. This signal by way of a lead 73, an armature 75, and a lead 77 is applied across potentiometer 52 Where the signal is modified by the position of its wiper. This position is controlled by motor 60 and signal which corresponds to mg is applied to potentiometer 53 where it is modified by a function the right hand portion of Equation 11 above.

Potentiometer 54 alone supplies the signal corresponding to W/S, the left hand portion of Equation 11, as the position of the wiper of potentiometer 54 is varied in accordance with'the weight of the aircraft; the surface being substantially constant. Norm-ally, the signal in lead 83 from potentiometer 5 4 is equal and opposite to the signal in lead 81 from potentiometer 53; consequently, no resulting signal appears at lead 85.

When a changefrom normal conditions occurs, the

balance between the signals in leads 81 and 83 is upset and a signal appears at lead 8 5. This signal is modified by the position of the wiper of potentiometer 55 and applied to a conventional amplifier 87. Potentiometer 55 stabilizes the system by preventing motor 60 from hunting because of its operating too rapidly for large signal changes due to large fluctuations in airspeed.

' The output of amplifier 87 energizes the variable phase winding 90 of induction motor 60 whose fixed phase winding 93 is constantly energized by a suitable source of alternating current. This motor is essentially a null seeking motor. In response to a signal input to amplifier 87, motor 60, by way of shaft 89 positions the wiper of potentiometer 52 to cancel out the signal. The position of this wiper, it will be remembered, corresponds to the computed angle of attack a It is obvious that, if desired, indicators with index and indicia such as at 97, 98, and 99 may be provided to give visual indications of the slope of the lift curve, the weight, and the angle of attack, respectively.' It is further obvious that bufier amplifiers as desired may be provided between the various potentiometers to prevent overloading of the otentiometers.

In angle of attack computer 26, the signal to amplifier 87 is normally zero. However, when a change in airspeed occurs, the dynamic pressure changes. The relativevalue of the impact pressure in tube 67 of airspeed sensor 62 with respect to the static pressure in tube 65 changes; the corresponding movement of diaphragm 63 moves shaft 71, moving the wiper of potentiometer 5 1 and changing the signal in lead 73. As a result, the signal at lead 81 no longer balances the signal at lead 83. A net signal corresponding to the change appears at lead 85. After amplification by amplifier 87, this signal energizes motor 60 which then positions the wiper ofpotentiometer 52 to cancel out the signal. This again brings about a condition where the combined signal from potentiometers 51, 52 and 53 equals the signal from potentiometer 54 and no resultant signal appears at amplifier 87; consequently, themotor stops and does not operate again until another change occurs. A' similar ac tion occurs when the position of the wiper of potentiometer 54 isva-nied because of a change in weight of the craft such as. may be caused by the consumption of fuel. The new angle of attack will be visually indicated by dial 99 which is moved by motor 60. Dial 99 may be calibrated to indicate the total angle of attack which may be written algebraically as in Equation 12. This motor also moves rotor 9-1 of an inductive. device 92 to develop a signal in the pitch channel for servomotor 10. In-

corresponds to the angle of attack or a The output ductivedevice 92 maybe arranged relative to potentiometer 52 to provide a fixed voltage corresponding to the angle of attack for zero lift when the angle of attack measured from zero lift a is zero. To the foregoing signal from inductive device 92 of the angle of attack computer is added the signal from the novel flight angle computer 28.

The flight path angle of the aircraft may be defined as the angle whose tangent is equal to the ratio of the vertical airspeed to the horizontal airspeed. In the embodiment herein, indicated airspeed is used for simplicity. For some flight conditions the difference between the indicated and true airspeed is sogreat that true airspeed must be used. The novel flight path computer is comprised of a potentiometer 120, a potentiometer 1'22, and a resolver 124. Potentiometer 120 whose wiper is positioned by shaft 71 develops an output corresponding to the airspeed. Potentiometer 122 whose wiper is connected to the shaft of a rate of climb indicator 126 develops an output corresponding to the rate of climb. The rate of climb indicator may be of a conventional type, moving shaft 127 an amount corresponding to the rate of climb. These voltages are fed to resolver 124.

Resolver 124 is an inductive signal generating device consisting of a single phase rotor 125 and a two phase stator. The two phase windings 128 and 130 of the stator are wound ninety electrical degrees apart. The output of potentiometer 122 corresponding to the rate of climb H, is applied to stator winding "12s and the output of potentiometer 120 corresponding to the airspeed V is applied to stator winding 130. The resultant of the field produced corresponds to the tan H V. When the windings of rotor 125 are not parallel to the resultant field an output corresponding to the displacement is developed. This is shown in Figure 2. This output is applied through amplifier 133 to the variable phase winding 134 on induction motor 136 whose fixed phase winding 138 is constantly energized by a suitable source of alternating current. The motor in response to energizetion drives the rotor to its null position. .The null position corresponds to tan- H/ V.

Motor 136 also displaces the rotor 141 of an inductive device 143 to developa signal in the pitch channel for the elevator. servomotor 10. At the same time, the motor may also displace an indicator 145 to give a visual indication of the flight path angle.

With the aircraft at an equilibrium condition at a pre scribed pitch attitude and the elevator surfaces at their normal positions, the signals from the pick-offs of vertical gyro 24, the angle of attack computer 26, and the flight path angle computer 28 are balanced, and the signal from follow-up device 16 is zeroso no net signal appears at the input of amplifier 30. When a change in external conditions occurs, as for example, a change in airspeed, the angle of attack computer 26 determines the angle of attack necessary for sustained level flight and develops a corresponding signal as motor 60 moves rotor 91 of inductive device 92. Atthe same time the flight path angle computer 28 determines the flight path angle of the craft and develops a corresponding signal 7 as motor 136 displaces the rotor 141 of the inductive device 143. These algebraically combined signals are applied to amplifier 30.

' Motor 10 in response to the output of amplifier 30 displaces the elevators 14 until the signal developed in follow-up device 16 is equal and opposite to the algebraic sum of the signals from computers 26 and 28 and the pick-off 23 of vertical gyro 24. The net input signal to amplifier 30 at this time is zero, so the motor stops. The displaced elevator surfaces place the craft in a different pitch attitude, however, and an error signal develops at inductive device 23 on vertical gyro 24. As

the pitch attitude signal becomes equal to the computer signals, the follow-up signal gradually operates servomotor 10 to return the elevator surfaces to a position to maintain a pitch attitude substantially as called for by the computers.

If the elevator displacement required to maintain the pitch attitude called for by the computers is substantial, it may be desirable to wash out persistant follow-up signals after a predetermined time delay so that the follow-up signal is zero as described in copending application Serial No. 217,988, filed March 28, 1951, by John C. Owen, now U. S. Patent No. 2,851,645, and assigned 'to the same assignee as the present application.

In order that the human pilot may control the airspeed and the rate of climb two identical command stations are provided: station 2% for the airspeed and station 261 for the rate of climb.

At station 2%, a manually operable knob 203 through shaft 202 turns a gear 2%, the wiper of a non-linear potentiometer 2115, the wiper of a potentiometer 207, and the rotor 299 of an inductive device 211. Potentiometer 205 by way of lead 213 supplies a signal for operating motor 214 through differential 216A to control the position of the throttle 216 for the power plant of the craft. Potentiometer 297 by way of lead 213, armature 75, and lead 77 provides a control signal for the angle of attack computer 26 when armature 75 is in the position shown in the drawing.

inductive device 211 has its stator 220 connected in parallel with stator 223 of an inductive device 225. The rotor 227 of inductive device 225 is fixed to shaft 71 which is movable by the dynamic airspeed sensor 62. When an error exists in the relative positions of rotors 2th? and 227 showing the actual airspeed of the craft is different from the desired airspeed, the corresponding signal developed is fed by leads 229, 231 and 233 to motor 235 when armatures 247 and 248 are in the position shown in the drawing to control the throttle to maintain the airspeed of the craft constant.

Gear 204 on shaft 2 32 turns a gear 250 whichis rotatable on shaft '71. An electromechanical interlock 251, however, prevents gear 256 and, consequently, gear 204 and 2% from being 'moved until after a switch 253 has been depressed. Moving switch 253 to a closed circuit position and moving knob 203 to anew position to change the airspeed of the craft causes a post 259 of a switch 26% on gear 25am move out of a recess 261 in a cam 262 which is suitably keyed to shaft 71. This com pletes a circuit from a direct current source such as battery 263 through a solenoid relay 267 to ground 269, energizing the relay.

T he energization of the relay pulls its armatures downwardly from the position shown in the drawing with the following results. The engagement of armature 270 with contact 271 form a holding circuit for maintaining relay 267 energized even though push button 253is released. The disengagement of armature 247 from contact 273 interrupts the control of motor 235 'by inductive devices 211 and 225. The engagement of armature 274 with contact 275 energizes a magnetic clutch 276 so that motor 214 can position throttle 216. The disengagement of armature 75 from contact 278 and its engagement with contact 2719 changes the control of the angle of attack computer 26 from the manual control potentiometer 207 to the airspeed controlled potentiometer 51.

Command station 201 is similar to command station 200. Depressing a push button 300 to release an electromechanical interlock 3ii1 and tuning a knob 303 then to a desired rate of climb of the craft turns a shaft 304, a gear 305 and the wiper of a potentiometer 306. As gear 305 turns a gear 307 which is rotatable on shaft 127, a post 311 of a switch 312 leaves recess 313 in a cam 314 which is keyed to shaft 127and completes a circuit through a solenoid relay 315 for the flow of cur rent from direct current source 263 to a ground 317.

. When relay 315 is energized, the armatures move downwardly from the position shown. An armature 330 engaging with a contact 331 forms a holding circuit for maintaining relay 315 energized even though push button 300 is released. Armature 248 disengaging from a contact 333 interrupts the control of motor 235 by inductive devices 211and 225. An armature 336 engaging a contact 337 energizes magnetic clutch 276 so that potentiometer 306 and a potentiometer 340 can operate motor 214 to position the throttle, the position of the wiper of potentiometer 340 corresponding to the weight of the craft. An armature 338 disengaging from a contact 339 and engaging with a contact 341 changes the control of the flight path angle computer 28 from that by potentiometers 306 and 340 to that by rate of climb controlled potentiometer 122.

When a new airspeed is desired without departing from level flight, button 253 is depressed and knob 203 is turned to the new airspeed, gear 204 turns gear 250 relative to shaft 71; and switch post 259 moves out of recess 261, completing a circuit between battery 263 and ground 269 through relay 267. The resulting energizetion of relay 267 pulls armatures 270, 75, 247 and 274 downwardly from the position shown in the drawing (the position assumed when relay 267 is deenergized). The engagement of armature 270 and contact 271 maintains the energization of relay 267. The disengagement of armature 247 from contact 273, interrupts the control of motor 235 by the ouput of inductive devices 211 and 225. 'The engagement of armature 274 with contact 275 energizes magnetic clutch 276 so that servomotor 214 engages with throttle 216 at the existing throttle position. The output of a potentiometer 205 whose wiper is positioned by command knob 203 and the ouput of potentiometer 340 whose wiper is positioned by weight knob 58 energizes servo 214 to move the throttle to a new setting.

The new throttle setting alone would cause the craft to ascend or descend as well as produce a forward acceleration or deceleration. The novel angle of attack computer 26 compensates for this tendency by developing a signal at inductive device 92 for the pitch channel so that the correct angle of attack will be maintained for sustained level flight. This sequency is based on the ability of an aircraft to assume a new pitch attitude quicker than it can attain a new airspeed.

As the airspeed increases, airspeed sensor 62 will cause shaft 71 to turn until, as the command value is reached, the recess 261 and post 259 are in registry. The circuit from battery 263 to ground 269 is opened and relay 267 is deenergized so its armatures will assume the positions shown in the drawing. The engagement of armature 247 with contact 273 permits the signal from inductive devices 211 and 225 to operate servomotor 235 to maintain this airspeed constant. Armature 75 engaging with contact 278 then applies a constant signal from potentiometer 207 to the angle of attack computer to maintain a constant pitch attitude since the angle of attack does not need to be changed unless a change in airspeed occurs.

When it is desired to change the rate of climb Without changing forward speed, button 300 is depressed and knob 303 moved to the new rate of climb. This displaces gear 307 relative to cam 314. Post 311 moves out of recess 213, closing a circuit from battery 263 to ground 317 and energizing relay 315. Armatures 330, 248, 336 and 338 move downwardly from the position shown. The engagement of armature 336 and contact 337 energizes magnetic clutches 276 thereby engaging servomotor 214 for operation of throttle 216 at its instant position. The engagement of armature 338 and contact 341 changes the operation of the flight path angle computer 28 from potentiometer 306 to 122. The disengagement of armature 248 from contact 333 interrupts the circuit from inductive devices 211 and 225 to motor 235. As the wiper of potentiometer 306 is displaced, the

resulting signal added to that of potentiometers 340 and 205 energizes servomotor 214 to move the throttles.

The power plant of the craft with the throttle at its position at the instant of engagement has been supplying the power required for the former flight condition. The

additional increment of throttle necessary for the new flight condition is supplied by motor 214.

'In the flight path anglecomputer 28, the output of otentiometers and 122 which, respectively, correspond to the airspeed and rate of climb are applied to resolver 124 whose output drives a motor 136. This motor moves the rotor of the resolver to a null position, positions an indicator .145 which shows the flight path angle of the craft with respect to the horizon, and displaces rotor winding 141 of inductive device 143- which develops a pitch control signal for servomotor 10. This displacement of the elevators by motor 10 places the craft in the proper pitch attitude and at the proper angle of attack to maintain the rate of climb set at knob 303. The pitch attitude is changed smoothly in a manner com patible with the angle of attack and the flight path angle.

As in the case of airspeed control, when the actual rate of climb attained corresponds with the rate of climb set at knob 303, post 311 will be in registry with recess 313;relay 315 will be deenergized, and its armatures will move to the position shown in the drawing. At this time the output of inductive devices 211 and 225 will be applied to servomotor 235 to maintain the aircraft at the fixed airspeed.

Thus, the action of the novel manual controller is eifective only when operated by the human pilot. When a new airspeed or rate of climb is attained, the holding circuit formed by armatures 270 and 330 will open the relay circuits. Controller knobs 203' and 303 must be reset by the operator for additional changes. This arrangement avoids the possibility of the throttle position changing in normal flight if the actual airspeed or rate of climb should differ momentarily from the command values.

The foregoing has presented a novel automatic pilot system in which the airspeed can be changed as desired by merely turning a control knob and yet the craft will maintain level flight. On the other hand by merely turning another knob, the craft may be placed in a desired rate of climb and yet maintain a constant airspeed. At the same time a visual indication may be made of the angle of attack and the flight path angle.

Although but one embodiment of the invention has been illustrated and described in detail, it is to be expressly understood that the invention is not limited thereto. Various changes can be made in the design and arrangement of the parts without departing from the spirit and scope of the invention as will now be understood by those skilled in the art.

' I claim:

1. An automatic pilot system for an aircraft, having a movable surface controlling the craft about its pitch axis, a servomotor for moving said surface, a first reference means for developing a signal corresponding to the deviation of said craft from a predetermined pitch attitude, second reference means for developing a signal corresponding to the angle of attack needed to sustain level flight, a third reference means responsive to rate of climb and airspeed of the craft for developing a signal corresponding to a desired flight path angle, and means connecting all of said reference means with said servomotor for operating the latter.

2. An automatic pilot system for an aircraft having a movable surface for controlling the craft about its pitch axis, a servomotor for moving said surface, a first reference means for developing a signal corresponding to the deivation of said craft from a predetermined pitch attitude, a second reference means responsive to rate of climb and airspeed of. the craft for developing a signal corresponding to a desired flight path angle, and means connecting said reference means with said servomotor for operating the latter.

3. An automatic pilot system for an aircraft having a movable surface thereon for controlling the craft about its pitch axis, servomotor means for moving said surface, reference means including means for measuring the rate of climb of said craft, means for measuring the airspeed of said craft, means responsive to both said measuring means for developing a signal corresponding to the angle whose tangent is the ratio. of the rate of climb to the horizontal airspeed, and means connecting said reference means and said servomotor for operating the latter by said signal.

4. An automatic pilot system for an aircraft having a movable surface for controlling the craft about its pitch axis, servomotor means for moving said surface, reference means for developing a control signal corresponding to the angle of attack needed to maintain level flight including first means for developing a signal corresponding to the airspeed of the craft, means connected with said first means for modifying said signal by a function of the angle of attack of the craft and by a fmnction of the slope of a curve representing the coeflicient of lift versus the angle of attack, second means for developing a signal corresponding to the weight per area of lift developing surface, means connected with said modifying means and said second means for comparing said last named signal and said modified signal to equate the two signals by changing the angle of attack function, said last named means actuating a signal device to develop said control signal, and means operatively connecting said reference means and said servomotor for operating the latter from said control signal.

5. An automatic pilot system for an aircraft having a movable surface thereon for controlling the craft about its pitch axis, servomotor means for moving said surface, reference means including means for developing a signal corresponding to the airspeed of the craft and for modifying said signal as a function of the angle of attack and the slope of the coeflicient of lift versus the angle of attack curve, means for developing a signal corresponding to the ratio of the weight of the craft to the lift developing surface, means for changing said angle of attack function to equate the modified signal and the signal which corresponds to the weight of the craft, said last named means also developing a control signal corresponding to the angle of attack, and means for connecting said last named means and said servomoto-r for operating the latter by said control signal.

6. An automatic steering system for an aircraft having a movable pitch control surface and a movable thrust control means, comprising pitch servo means for operating the pitch control surfaces, thrust servo means for operating the thrust control means, manually operable command means for ordering a predetermined airspeed, means operatively connecting said command means: and said thrust servo means for actuating the latter an amount sufficient to maintain said predetermined airspeed and means operatively connecting said command means and said pitch servo means and responsive to said airspeed for automatically modifying the angle of attack of the craft to keep said craft from ascending and descending.

7. An automatic steering system for an aircraft having a movable pitch control surface and a movable thrust control means, comprising pitch servo means for operating said pitch control surfaces, thrust servo means for operating said thrust control means, command means for ordering a predetermined rate of climb of the craft, means operatively connecting said command means and said thrust servo means for actuating the latter an amount sufficient to maintain an existing airspeed, means responsive to said airspeed and operatively connected with said pitch servo mechanism for operating the latter to maintain a predetermined angle of attack of the craft, and means operatively connected with said command means t 10 and also connected with said pitch servo mechanism to place the craft in a pitch attitude to maintain a prescribed flight path angle whereby said last two means maintain said craft at said predetermined rate of climb. I

8. An automatic pilot system for an aircraft having a movable thrust control means thereon, comprising first means for automatically moving said thrust control for maintaining a constant airspeed for the craft, control means for presetting an airspeed, second means operatively connected with said control means for moving said thrust control to cause said craft to reach said preset airspeed, and means operatively connected with said control means and said first and second means for rendering said first means ineffective and said second means efiective on said thrust control when said preset airspeed is dilferent from the existing airspeed and when the airspeed reaches said preset airspeed for rendering said second means: ineffective and said first means eliective to maintain said preset airspeed.

9. An automatic pilot system for an aircraft having a pitch attitude maintaining means and a movable thrust control means thereon, first means connected with the thrust control means for automatically moving the latter to maintain the airspeed of the craft at a constant value, control means for presetting an airspeed, second means operatively connected with said control means and said thrust control means for moving the latter to cause said craft to reach said preset airspeed, means operatively connecting said control means, and said first and second means for rendering said first means ineffective and said second means effective on said thrust control when said preset airspeed is different from the existing airspeed and when the airspeed reaches said preset airspeed for rendering said second means ineffective and said first means effective to maintain said preset airspeed, and angle of attack means connected with the pitch attitude maintaining means responsive to said airspeed for operating said attitude maintaining means to keep the craft from ascending or descending due to the change in airspeed.

10. An automatic pilot system for an aircraft having a movable thrust control means thereon, first means for automatically moving said thrust control for maintaining the airspeed of the craft constant, control means for presetting an airspeed, second means operatively connected with said control means and said thrust control for moving the latter to cause said craft to reach said preset airspeed, means operatively connected with said control means and said first and second means for rendering said first means ineffective and said second means effective on said thrust control when said preset airspeed is different from the existing airspeed and for rendering said second means ineflective and said first means efiective when the airspeed reaches said preset airspeed, and means connected with said control means and responsive to the airspeed of said craft for maintaining the craft at a flight path angle to keep the craft from ascending or descending due to the change in airspeed.

11. An automatic pilot system for an aircraft, having a movable pitch control surface and a movable thrust control means thereon, first means for automatically moving said thrust control for maintaining a constant airspeed for the craft, control means for presetting an airspeed, second means operatively connected with said control means for moving said thrust control to cause said craft to reachsaid preset airspeed, means operatively connected with said control means and said first and second means for rendering said first means ineffective and said second means effective on said thrust control when said preset airspeed is different from the existing airspeed and for rendering said first means effective and said second means ineffective when the airspeed reaches said preset airspeed and angle of attack means responsive to said airspeed for maintaining said craft at a correct angle of attack to keep the craft from ascending or descending due to the change in airspeed, flight path angle comput- -11 ing means responsive to the change in airspeed for maintaining the craft at a predetermined rate of climb, servo means for moving said pitch control surface and means connecting said angle of attack means, said flight path angle computing means and said servo means for operating the latter.

12. In a system for controlling the attitude of an aircraft, means for developing a signal corresponding to the airspeed of the craft, means for developing a signal corresponding to the rate of climb of the craft, means connected with said signal developing means for receiving said signals and developing an output corresponding to the angle whose tangent is the ratio of the rate of climb signal to the airspeed signal 13. In a system for controlling. the attitude of an aircraft, first means for developing a signal corresponding to the airspeed of the craft, means connected with said first means for modifying said signal by a function of the angle of attack of the craft and by a function of the slope of a curve representing the coefiicien-t of lift versus the angle of attack, second means for developing a signal corresponding to the weight per area of lift developing surface, means connected with said modifying means and said second means for comparing said last named signal and said modified signal to equate the two signals by changing the angle of attack function, said last named means actuating a signal device to develop a control signal.

12 14. An automatic pilot system for an aircraft having a movable thrust control means and movable pitch control means thereon, means for automatically moving said thrust control for maintaining the airspeed of the craft constant, first means for maintaining said craft in a predetermined rate of climb, control means for presetting I a rate-ofclimb of the aircraft, means operatively connected with said control means and said pitch control means for moving the latter to cause said craft to attain said rate of climb, means operatively connected with said control means and said first and second means for rendering said first means inelfective and said second means effective on said pitch control means when said rate of climb is difierent from the existing rate of climb and means for rendering said second means inefiective and said first means eflective when the rate of climb reaches said preset rate of climb.

References Cited in the file of this patent UNITED STATES PATENTS 2,415,092 Frische Feb. 4, 1947 2,620,149 Strother Dec. 2, 1952 2,701,1ll Schuck Feb. 1, 1955 2,798,682 Alderson et al. July 9, 1957 FOREIGN PATENTS 495,981 Belgium Sept. 16, 1950 

